The present invention relates to the cooling of turbine rotor blades and stationary vanes (both of which are generically referred to herein as xe2x80x9cturbine bladesxe2x80x9d unless otherwise indicated). The invention relates more particularly to cooling of turbine blades using a coolant supplied to internal passages in the blades.
A turbine produces rotational power by receiving high-temperature, high-pressure gases such as combustion gases from a fuel combustor, and expanding the gases to a lower temperature and lower pressure via an alternating series of stationary vanes and rotating blades. A gas turbine may have a single xe2x80x9cstagexe2x80x9d consisting of a row of stationary vanes followed by a row of rotor blades, or it may have two or more such stages in series. In high-performance gas turbine engines, the temperature of the combustion gases entering the first stage of the turbine typically is so high that the available materials for constructing the stationary vanes and rotor blades are not capable of withstanding the extreme temperature without some type of active cooling of the blades and vanes. Thus, modern advances in gas turbine technology have largely been made through discoveries of improved materials capable of withstanding higher temperatures, coupled with improved cooling schemes.
The efficiency of gas turbines generally goes hand-in-hand with the turbine inlet temperature, such that higher turbine inlet temperatures provide higher efficiencies in general. These higher temperatures increase the challenge of cooling the blades and vanes adequately. Conversely, anything that reduces the effective temperature of the hot gases passing through the turbine without doing a corresponding amount of work results in a reduction in turbine efficiency. This leads to a tradeoff in conventional gas turbines because the blades are typically cooled by film cooling techniques. In film cooling, a cooling fluid (typically air in most gas turbines) is supplied through internal passages formed in the turbine blade and is ejected from the passages through holes in the outer surface of the blade, such that the cooling fluid flows over the outer surface to be cooled and forms a protective layer of fluid that is substantially cooler than the hot gases passing through the turbine, thus effectively insulating the blade surface against the hot gases. It is typical to have a relatively large number of film cooling holes around the leading edges of turbine blades, especially in the first stage or first few stages where the temperatures of the hot gases are greatest, and to have additional film cooling holes distributed over the suction-side and pressure-side surfaces of the blades, and perhaps film cooling slots in the trailing edges of some blades.
It will be appreciated that film cooling thus involves injecting cooling fluid in the main gas flow path of the turbine, which reduces the effective temperature of the gases passing through the turbine. This leads to a reduction in the efficiency of the turbine. In extreme cases, the total mass flow of cooling fluid ejected through film cooling holes may represent 20 percent of the mass flow of the hot gases, or more, leading to efficiency reductions of 10 percent or more.
Accordingly, it would be desirable to improve the cooling of turbine blades, enabling higher turbine inlet temperatures and correspondingly improved turbine efficiencies for a given type of blade material.
The present invention addresses the above needs and achieves other advantages, by providing a cooled turbine blade (either a rotor blade or a stationary vane) that comprises a blade structural member whose primary function is to withstand the various loads exerted on the blade and maintain structural integrity of the blade, and a heat-transfer sheath that surrounds the outer surface of the structural member. A plurality of coolant passages are formed between the structural member and the heat-transfer sheath. Thus, when coolant is passed through the coolant passages, the heat transferred to the sheath from the hot gases is in turn transferred to the coolant, which is then removed from the blade, thus cooling the blade.
In preferred embodiments of the invention, the coolant passages are closed, such that they do not emit any coolant into the main gas flow path of the turbine. In these embodiments, the coolant passages in the blade are in fluid communication with coolant supply and exhaust manifolds formed, for example, in the disk supporting a rotor blade or in one of the shrouds of a stationary vane. Each coolant passage is a closed loop such that all coolant that flows through the passage into the blade subsequently flows back out of the blade and is recovered, with the possible exception of very small amounts of coolant leakage that may occur, for example, at sealed connections between a rotor blade and its disk or between a stationary vane shroud and the casing in which it is mounted. Thus, substantially no coolant is dumped into the main gas flow path of the turbine, thereby improving potential turbine efficiency.
The coolant passages can be formed in the outer surface of the blade structural member, such as by machining the outer surface. Alternatively, the channels can be machined or otherwise formed in the inner surface of the sheath. Conveniently, the passages can be machined as channels of rectangular or square cross-section; bonding the heat-transfer sheath onto the outer surface of the structural member then closes the channels to form closed passages.
In preferred embodiments of the invention, the coolant supplied to the coolant passages comprises liquid water. As the water flows through the passages, heat transfer into the water from the sheath causes steam to be formed. The coolant may exit the passages primarily in the form of saturated steam. In order to maintain the walls of the passages bathed with liquid water as much as possible so that the desired high heat transfer rate into the coolant is maintained, it is preferred to size the passages so that surface tension of the water keeps the water adhered to the passage walls. This can be accomplished by configuring each passage in cross-section as a parallelepiped (e.g., a rectangle or square) each edge of which is about 0.5 to 1.3 mm (0.02 to 0.05 inch) in length.
The heat-transfer sheath can comprise various materials preferably of high thermal conductivity. Examples of suitable materials include but are not limited to copper, nickel, alloys such as Narloy-Z (a high-strength copper alloy). The sheath can be attached to the blade structural member in various ways, with diffusion bonding being the preferred technique. The sheath preferably is formed in multiple separate pieces that collectively cover the structural member. The sheath preferably is relatively thin, for example, about 1 to 2 mm (0.04 to 0.08 inch).
The invention also may enable damping of blade vibrations to be accomplished by fluid damping from the coolant in the internal coolant passages, as opposed to the use of external damping devices often used in conventional turbines. More particularly, frictional damping devices that rub against adjacent surfaces during blade vibrations are frequently used in conventional turbines in order to reduce the magnitude of blade vibrations to acceptable levels so that the blades have adequate fatigue life. Frictional dampers, being external to the blades, tend to disturb the blade aerodynamics, which leads to reduced turbine efficiency. Such dampers also are subject to wear that can reduce their effectiveness and eventually may necessitate their replacement. Frictional dampers also represent additional parts that must be manufactured, inventoried, installed, monitored, and replaced when needed. If a damper should fail and break loose during turbine operation, it could cause damage to the turbine and/or to components downstream of the turbine.
In contrast, the fluid damping provided by the coolant, such as liquid water, flowing through the coolant passages between the sheath and blade structural member of the present invention requires no extra parts and hence no additional cost, does not disturb the blade aerodynamics, and does not employ components that could break loose and cause damage. The fluid damping is essentially out of phase with primary bending and shear stresses in the blade, such that the damping can reduce internal shear forces and deflections.